Conventionally, a rotorcraft rotor comprises a hub driven in rotation about an axis of revolution by an outlet shaft of a power transmission main gearbox, also referred to as a drive shaft, together with at least three blades fastened to the hub via appropriate hinges, in particular via respective laminated spherical abutments dedicated to each of the blades.
On the assumption that each blade is fixed relative to the hub, it should be recalled that the rotor constituted in this way is a rigid rotor. While hovering, the distribution of aerodynamic forces along a blade gives rise to a bending moment distribution presenting a very large value at the root of the blade. When flying in translation, the blade that is said to be “advancing” carries more than the blade that is said to be “retreating” because of the unequal air speeds as described in greater detail below.
Consequently, the resultant of the aerodynamic forces exerted on a blade does not have the same value in each direction, nor even the same point of application: the fixed-end or “restrained” bending moments at the root of the blade is thus high and variable, thereby generating alternating mechanical stresses leading to a fatigue phenomenon that is harmful for the materials. Furthermore, the resultant of the aerodynamic forces from all of the blades is no longer carried by the axis of the rotor, thereby creating a roll moment that increases with increasing speed and that can make it difficult to balance forces while flying in translation.
In order to remedy those drawbacks, it is known to hinge the blades to the hub about respective axes that are perpendicular to the drive shaft and referred to as the vertical flapping axes that correspond to a vertical flapping hinge capable of transferring a force of arbitrary orientation but not capable under any circumstances of transferring a moment. Consequently, if the blade is hinged to the hub, the bending moment where it is attached is zero. To achieve blade equilibrium, centrifugal forces keep the blade in place after it has risen a certain amount, thereby producing a cone shape of angle a0.
Under such conditions, there is no longer a large amount of roll moment when flying in translation and the blades no longer revolve in a plane, rather their outer ends describe a very flat cone. In practice, the flapping axis is then not on the axis of revolution but is offset therefrom by a distance a, referred to as the eccentricity.
It should also be recalled that in order to provide a helicopter with lift in its various different configurations, it is necessary to be able to control the lift provided by the rotor and to vary it. That is why a pitch hinge is provided of axis that is substantially parallel to the span of the corresponding blade. This new degree of freedom enables the lift of the blade to be controlled by acting on a general pitch control, and also enables pitch to be caused to vary cyclically, thus serving to control the plane of rotation of the blades, which blades then describe a cone of axis that no longer coincides with the drive axis: the resultant of the forces applied to the hub changes direction whenever the plane of the rotor changes direction. This gives rise to moments about the center of gravity of the helicopter, thus enabling it to be piloted.
As mentioned above, the plane of rotation of the blades may be different from a plane that is perpendicular to the drive shaft. Under such conditions, it is necessary to hinge each blade in drag since the end of each blade is at a variable distance from the rotor shaft. Otherwise, inertial forces would necessarily appear, giving rise to alternating bending moments in the plane of each blade. Such a drag hinge is provided by hinging the blade about a drag axis that is substantially parallel to the axis of the rotor, and consequently substantially perpendicular to drag forces. To enable such a blade to be driven from the drive shaft, it is naturally necessary for the drag hinge to be far enough away from the rotor axis to ensure that the moment due to centrifugal forces balances the moment due to drag and inertial forces, thus requiring the drag axis to be offset by an eccentricity e, and with this requiring that the “drag” angle δ is not too great.
Consequently, the blades of a hinge rotor of a rotary wing aircraft, in particular a helicopter, can describe the following four movements:
i) rotation about the axis of the rotor;
ii) vertical rotation about the flapping axis, made possible by the vertical flapping hinge;
iii) horizontal rotation about the drag axis, also referred to as the horizontal flapping axis, made possible by the horizontal flapping hinge or drag hinge; and
iv) rotation about the axis of the blade made possible by a pitch hinge (not specific to hinged rotors).
As described in patent FR 2 497 073, for example, the above three rotations ii), iii), and iv) may be made possible by a single member such as a laminated spherical abutment.
Nevertheless, the oscillations of each blade about its drag axis can become coupled in unstable manner with the elastic deformation modes or movements of the airframe, in particular oscillations of the helicopter while standing on the ground via its landing gear: this is the origin of the so-called “ground resonance” phenomenon that can be dangerous for the aircraft when the resonant frequency of the oscillation of the blades about their drag axis is close to one of the resonant frequencies of the oscillations of the aircraft relative to its landing gear.
The remedies for that phenomenon consist in introducing damping on the drag axes by means of a damper type device.
Such dampers have resilient return means of determined damping and stiffness for combating resonance phenomena, in particular ground resonance phenomena and also drive system resonance phenomena, that can appear in particular on helicopters.
When the blades of the rotor are excited in drag, the blades move away from their equilibrium positions and may become unevenly distributed in the circumferential direction, thereby creating an unbalance by moving the center of gravity of the rotor away from its axis of rotation. Furthermore, blades that are spaced apart from their equilibrium positions oscillate about said positions at a frequency ωδ, which is the resonant frequency of the blades in drag, also referred to as the first drag mode or as the resonant mode in drag.
If Ω is the frequency of rotation of the rotor, it is known that the fuselage of the helicopter is thus excited at the frequencies |Ω±ωδ|.
When standing on the ground via its landing gear, the fuselage of the helicopter constitutes a mass-spring-damper system comprising a mass suspended above the ground by a spring, and a damper in each undercarriage. With the fuselage resting on its landing gear, there are therefore resonant modes in vibration in roll and in pitching. There is thus a risk of ground instability when the excitation frequency |Ω+ωδ| or |Ω−ωδ| of the fuselage on its landing gear is close to the resonant frequency of oscillation, which corresponds to the so-called ground resonance phenomenon. To avoid instability, it is known to begin by seeking to avoid any occasion on which these frequencies cross, and if such crossing cannot be avoided, then the fuselage on its undercarriage must be damped sufficiently, as must the blades of the main rotor in their drag movements.
Consequently, the stiffness of the drag dampers of the blades of a main rotor needs to be selected so that the resonant frequency of the blades in drag lies outside a potential ground resonance zone, while simultaneously also providing sufficient damping, since when the speed of rotation of the rotor passes through the critical speed, both when speeding up and when slowing down, the movements of the blades must be damped sufficiently to avoid entering into resonance.
That is why drag dampers having resilient return means of determined stiffness are also referred to as frequency adapters.
In general, the stiffness of the damper determines an equivalent angular stiffness that opposes angular flapping of the blades relative to the hub about their drag axes. It is thus possible to increase the frequency of the resonant mode of the blades in drag in order to move that frequency away from the two above-mentioned resonance phenomena.
The equivalent angular stiffness is proportional to the square of the lever arm between the damper and the drag axis of the blade, i.e. the distance between the drag axis and the axis passing through the centers of the two ball joints of the damper, where such ball joints are necessary in this application.
Document FR 2 653 405 presents two different configurations for such dampers.
Thus, according to that document, the head of the rotor is provided with means including an annular central portion, an intermediate portion provided with one cavity per blade, and then a peripheral portion.
Each blade is then secured via its root to a sleeve that is fastened on a laminated spherical abutment arranged in one of said cavities.
Furthermore, in a first embodiment, one rotary damper per blade is secured to the peripheral portion of the hub. The rotary damper is then a resilient return member with incorporated damping due to shear in a viscoelastic material presenting very high deformation remanence.
In order to be able to damp the drag movement of a blade, the rotary damper is connected by a connecting rod to a sleeve of the blade. To guarantee good operation, each connecting rod is then substantially perpendicular to a direction passing via its point fastened to the fork of the associated rotary damper and the axis of revolution of said rotary damper.
That first embodiment is suitable for rotorcraft provided with a rotor having three blades.
Since the lever arm of that device is relatively small, it is appropriate to use a rotary damper that is overdimensioned and thus bulky.
Thus, for rotorcraft having a rotor with four or more blades, document FR 2 653 405 proposes a second embodiment.
In that second embodiment, a rotary damper is fastened inside each sleeve, the rotary damper of a blade being connected to an adjacent blade by a connecting rod.
Compared with a more conventional configuration in which the dampers are interposed between each blade and the hub of the rotor, the inter-blade configuration of a damper enables the lever arm to be increased between the dampers and the drag axes of the blades, but it also causes two dampers to participate with each blade in order to avoid ground resonance. The stiffness of each damper may be limited accordingly, and one resulting advantage is a lower level of static force introduced by mounting each damper as an inter-blade adapter. That configuration is therefore very favorable for combating ground resonance.
However, inter-blade mounting does not serve to damp overall drag movements and therefore requires drag abutments in order to avoid damage when the rotor is starting, and above all when it is braking.
Consequently, the prior art provides two distinct and alternative embodiments, namely:                arranging a rotary damper on the hub that is connected to the sleeve of a blade by means of a connecting rod; or        arranging a rotary damper in the sleeve of a blade, which rotary damper is connected by a connecting rod to an adjacent blade.        
Each embodiment then presents its own advantages and drawbacks.
Nevertheless, independently of the embodiment that is selected, coupling is observed between pitch and flapping. A flapping movement of the blade gives rise to a modification to its pitch. Likewise, a modification to the pitch of the blade under the control of the pitch control rod gives rise to a flapping movement of the blade.
This coupling between pitch and flapping arises mainly because of the particular positioning of the pitch control rod of the blade relative to the flapping and pitch axes of the blade.
Furthermore, it is observed that varying the pitch of the blade gives rise to excessive stress on the rotary damper.